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Full text of "Measurements of the Lateral and Directional Stability and Control Characteristics of a P-51H Airplane (AAF No. 44-64164)"

RESEARCH-MEMORANDUM 



Source of Acquisition 
CAS I Acquired 



for the 



Air Materiel Coramand, Army Air Forces 



MEASUREMENTS OF THE LATERAL AND DIRECTIONAL STABILITY 
AND CONTROL CHARACTERISTICS OF A $4$3M AIRPLANE 
(AAF No, kU-€kl&i) 

Christopher C. Kraft, Jr. and J. P. Reeder 

Langley Memorial Aeronautical Laboratory 
7^ Langley Field, Va. 




act 




LU- 



NATION AL advisory committee 

FOR AERONAUTICS FILE COi^Y 

Washington ft fc» return* tj 

JAN Z7 1948 «••«•• #f th« Nati 



CONFIDENTIAL 



•ourity Inform;itK»ri 



lifted tt* »9tkm\ 

fm AtrofttutiN 



NACA EM No . L7L1 




CONFIDENTIAL 



NATIONAL ADVISORY COMMITTEE FOE AERONAUTICS 



RESEARCH MEMORANDUM 
for the 

Air Materiel Command, Army Air Forces 



MEASUREMENTS OF TEE LATERAL AND DIRECTIONAL STABILITY 
AND CONTROL CHARACTERISTICS OF A P-51H AIRPLANE 
(AAF No. kk~6kl6h) 
By Christopher C. Kraft, Jr. and J. P. Reader 

SUMMARY v 



Flight tests of a F-51H airplane with two different vertical— tail 
assemblies were made to determine lateral and directional stability 
and control characteristics. The airplane had satisfactory directional 
stability in the landing, approach, and wave-off conditions with either 
tail. In the power— on clean and glide conditions, however 9 the airplane 
had weak directional stability with the original tail. The production 
tail, which had a 7— inch fin extension and a shorter span rudder, 
improved the directional staMlity in the power-on clean and glide 
conditions, but the stability was still weak in the power— on clean 
condition. Increased altitude in either case caused a slight decrease 
in the staMlity. The rudder-trim-force change with speed with either 
vertical— tail assembly was high. The general aileron control character- 
istics were satisfactory hut the aileron effectiveness failed to meet 
the Army handling-qualities requirements. 



INTRODUCTION 



At the request of the Air Materiel Command, Army Air Forces, flight 
tests have "been made to determine the lateral and directional staMlity 
and control characteristics of a North American F-51H airplane with the 
original and the production vertical— tail assemblies. Tests of the 
longitudinal stability and control and stalling characteristics will "be 
presented in a later paper. 

For previous tests- of the F-51D airplane, the NACA designed and 
built a fin extension for this airplane to improve its directional 
stability. Very satisfactory results were obtained from the addition 
of this extension as reported in reference 1. When the B-51H airplane 
was "brought out, the North American Aviation, Inc. conducted preliminary 




2 



CONFIDENTIAL 



NACA EM No. L7L11 



J * # ; increased directional stability. They therefore added a fin extension 
* in accordance with the design previously developed by the NACA for the 

• • *• P— 5 ID airplane. This extended tail on the B-51H airplane is designated 

• • as the production tail. At the time these tests vere "being made 9 the 

Army Air Forces were very much interested in obtaining this particular 
• • airplane for use in the war in the Pacific. Therefore, the NACA conducted 
tests on "both the original and the production tails in order to determine 
the improvement in directional stability that resulted from this change. 



DESCRIPTION OF AIRPLANE AND TESTS 



The is a low— wing fighter airplane equipped with a 

Msrlin V-I65O-9 engine (fig. l). The original vertical tail had a full- 
span rudder with -±30° travel and the production vertical tall had a 
7— inch fin extension and a shorter span rudder with the travel limited 
to ±25°. The original rudder had a O.k to 1 unbalancing tab which also 
served as a trim tab, whereas the production rudder had only a trim tab. 
Both fine were offset 1° left from the thrust axis. A photograph of 
the airplane showing a comparison of the two vertical tails is shown 
in figure 2 and a drawing of the two tails, in figure 3. Photographs of 
the airplane with the production tail are shown in figures k to 6. 
Pertinent dimensions of the airplane are given in table I. 

All of the control surfaces were metal covered^and the ailerons 
had sealed internal balances. The variations of the elevator and aileron 
position with stick position and rudder— angle variation with rudder- 
pedal position are shown in figures 7 to 9 the friction in these 
control systems is shown in figure 10. The stretch in the aileron 
control system was 2.5° per 20 pounds of control force, and the stretch 
in the rudder control system was 7.5° per 100 pounds of rudder— pedal 
force. These characteristics are shown in figure 11. 



The airplane was tested in the configurations shown in the following 
table. The center of gravity for these tests was at 2^.^ percent mean 
aerodynamic chord at take-off with the wheels down. 



Condition 


Power 


Flaps 


Gear 


Canopy 


Glide (power— off clean) 


Engine 


idling 


Up 


Up 


Closed 


Power on clean 


k6 in. 


Eg at 2700 rpm 


Up 


Up 


Closed 


Landing 


Engine 


idling 


Down 


Down 


Open 


Approach 


23 in. 


Hg at 2700 rpm 


Down 


Down 


Open 


Wave-off 


h6 in. 


Hg at 2700 rpm 


Down 


Down 


Open 



In addition, tests were run at high altitude in the glide and power- 
on clean conditions. These tests were run for both the vertical— tail 
assemblies . 

The sideslip data given in this paper were obtained by the steady- 
sideslip method; that is, the airplane was sideslipped to a certain angle, 

couFiDErrriAL 



NACA EM No. L7L11 



CONFIDENTIAL 



3 



and when conditions were steady 9 a record was taken of the various 
quantities indicated in the section on instrumentation. 

INSTRUMENTATION 



Standard NACA photographic recording instruments were used to o"btain 
the data contained in this paper along with pilot's readings of altitude, 
free— air temperature , and fuel— gage readings. The following recording 
instruments were installed in the test airplane: 

Accelerometer (three— component ) 

Stick— force recorder 

Rudder— pedal— force recorder 

Airspeed recorder 

Eoll turn meter 

Pitch turn meter 

Yaw turn meter 

Eecording inclinometer (angle of hank) 

Sideslip— angle recorder 

Control— posit on recorders 

Timer (synchronizing all records) 

The airspeed was measured "by means of a swivel ing static head and 
a shielded total head mounted on a "boom approximately 1 chord length 
ahead of the right wing tip. (See fig, 5.) The airspeed installation 
was calibrated for position error by means of a trailing airspeed head 
that measured the static pressure in the free— air stream. The trailing 
airspeed head and the reel were "built into an auxiliary droppable fuel 
tank which was attached to the airplane under the right wing and operated 
"by the pilot from the cockpit. With the airspeed head trailing, the 
airplane was flown at a series of speeds from the stall to 250 miles 
per hour and the position error of the airplane f s static pressure head 
was determined from these data. The term "calibrated airspeed" as used 
in this paper may he defined by the following equation: 

V c = 1*5.08 f Q /q£ 

where 

V c calibrated airspeed in miles per hour; that is, the reading in 

miles per hour that would be given by a standard Army— Navy 
airspeed meter if it were connected to a pitot static system 
free from position error 

f G standard sea— level compressibility correction factor 

q c pressure differential in inches of water between total and static 

head j corrected for position error 



CONFIDENTIAL 



CONFIDENTIAL 



NACA RM No. L7L11 



The jaw vane used to measure the angle of sideslip was mounted on 
a "boom approximately 1 chord length ahead of the left wing tip. (See 
fig. 5.) The term "indicated sideslip" used in this paper is the 
uncorrected value given by the yaw vane. The change in sideslip angle 
is "believed to he correct ^ hut the exact magnitude may he in slight 
error "because of the lack of symmetry of the yaw vane and/or angularity 
of flow at the yaw vane. 



LATERAL AM) DIRECTIONAL STABILITY AND CONTROL 



The discussion of results that follows is "based on the flying- 
qualities requirements set forth in reference 2. 



Dynamic Directional Stability 

The dynamic directional stability of the airplane was tested in all 
of the different configurations at various speeds. These tests were 
performed "by trimming the airplane at a certain speed and starting an 
oscillation "by deflecting and releasing the rudder quickly or by starting 
the oscillation from about 5° of steady sideslip. There was no oscil- 
lation of the rudder itself in any configuration at any speed or altitude. 
In the power— on clean condition,, when the oscillation was started from 
a left sideslip or a right rudder kick, the ensuing motion of the 
airplane resulted In a spiral to the left. Typical time histories of 
these oscillations are shown in figures 12 to 16. 



Static Directional Stability 

Power— off clean condition .— In the power-off clean condition 
(figs. 17 and l8) , the airplane with the original vertical tail possessed 

d&r 

positive rudder— fixed and rudder— free stability , but the parameters 

and — E (the variation of rudder angle and rudder force required for 
dp 

trim with sideslip angle) through zero sideslip were low; this was 
especially so at the speed of 150 miles per hour (fig. 17(a)). With the 
production tail the airplane had good rudder— free and rudder— fixed 
stability and showed a marked improvement over the airplane with the 
original tail assembly. The dihedral effect in both configurations was 
positive and satisfactory. 

The effect of altitude was to decrease the stability in both of the 
configurations (fig. 18) except at 150 miles per hour with the original 

<* & r / 

tail where there was little change in . (The actual values of — i 

dp d3 



CONFIDENTIAL 



NACA RM No. L7LL1 



CONFIDENTIAL 



5 



and - for the conditions tested are given in tabid II. These values 

dp 

are only approximate "because of the scatter of test points in some of 
the conditions tested.) 

Power— on clean condition .— The airplane in the power— on clean 
condition (figs. 19 and 20) had positive rudder— fixed and rudder— free 
stability through zero sideslip with "both the original and the production 
vertical— tail configurations. However, with the original tail, at the 
speed of 150 miles per hour (fig. 19(a)) a flattening of the rudder^- 
force variation occurred "between 7° and 16° of left sideslip angle, and 
the airplane had neutral to negative stick— fixed and stick— free dihedral 
effect "beyond 10° left sideslip and neutral stick— free dihedral effect 
"beyond 15° right sideslip. The large angles of sideslip that can "be 
obtained with this configuration are also oh Jectionable at this speed. 
At the higher speeds, 300 and kOO miles per hour, the directional 

d6-~ 

stability improved "but the values of — =- were still low. 

dp 

Although the airplane with the production tail was an improvement 
over the airplane with the original tail, the rudder— free stability 
through zero sideslip at 150 miles per hour was still low. There was 
no negative dihedral effect at any speed and the stability at high 
speeds was satisfactory. It should be noted that the increased tail 
height corrected the negative dihedral effects experienced at large 
angles of sideslip with the original tail. 

The effect of altitude was to decrease the stability in both 
configurations and at all the speeds tested. At 150 miles per hour 
with the original tail, the rudder— f orce curve reverses in slope 
from 10° to 20° left sideslip and reduces to 5 pounds of rudder force 
at 20° left sideslip. The negative dihedral effect at large sideslip 
angles was also pronounced at high altitude (fig. 20(a)). At both the 
high speeds tested, 300 and 350 miles per hour, the rudder— fixed stability 

was low and the slope appeared to be practically zero through the 

dp 

r angle of left sideslip angles. 

The airplane with the production tail still had positive rudder- 
fixed and rudder— free stability through zero sideslip at high altitude, 
but the rudder force flattened out between 6° and 20° left sideslip so 
that zero change in rudder force was required to Sideslip the airplane 
in this range of sideslip angles. There was a slight amount of negative 
stick— fixed dihedral effect beyond 12° left sideslip but this was not 
considered objectionable (fig. 20(b)). At the higher speeds tested at 
high altitude, both the rudders-fixed and rudder— free stability were 
satisfactory. 



CONFIDENTIAL 



CONFIDENTIAL 



NACA RM No. L7L11 



Power— off landing condition ,— In the powejx>ff landing condition 
(fig . 21), the airplane had satisfactory directional stability 
at 120 and 150 miles per hour for "both the original and production 
vertical— tail configurations. The production tail had higher values of 
the stability parameters and the rudder— force and rudder— angle variations 
with sideslip were more linear than the original tail. 

Approach condition .— In the approach condition (fig* 22) , the 
airplane had satisfactory directional stability at 120 and 150 miles 
per hour for both the original and production vertical— tail configurations. 
The original tail showed a slight nonlinearity in the variation of rudder 
force and rudder angle with sideslip which the production tail did 
not encounter. 

Wave— off condition .— In the wave-off condition (fig. 23) , the 
airplane had good directional stability with "both the original— and 
production— tail configurations at 120 miles per hour, the only speed 
tested in this condition. 



Rudder Control Power 

Adverse aileron yaw .— The adverse aileron yaw of the airplane 
(fig. 2k) was measured by performing aileron rolls with various amounts 
of aileron deflection with the rudder fixed and allowing the airplane 
to reach maximum sideslip angle. These data are plotted in figures 2k (a.) 
and 2k(\>) as the variation of« change in sideslip angle with change in 
total aileron deflection. These tests were made at 135 miles per hour 
in the power— off landing condition and at 150 miles per hour in the 
powej>-on clean condition. 

In the 135-mile— per— hour landing condition, 5— percent total aileron 
deflection produced approximately 1° of sideslip with both the original— 
and production— tail configurations which satisfies the requirements 
of reference 2. However, in the 150-mile— pei^-hour powei — on clean 
condition, in left aileron rolls, the amount of sideslip obtained was 
much more than 1° per 5 percent of the total aileron deflection. The 
airplane exhibited rapidly increasing sideslip angles in these maneuvers 
which would have resulted in dangerous attitudes in yaw if the pilot 
had permitted maximum sideslip angles to he reached. This unsatisfactory 
condition existed with either the original or production tail. 

Rudder control power to overcome adverse aileron yaw .— In order to 
measure the power of the rudder to overcome adverse aileron yaw (fig. 25) 
the airplane was rolled abruptly out of U5 banked turns using full 
aileron deflection and varying amounts of rudder deflection at 150 miles 
per hour in the power— on clean condition. The change in sideslip angle 
obtained hy application of varying amounts of change in rudder angle is 



C 0W WEWI1AL 



NACA EM No. L7L11 



CONFIDENTIAL 



7 



plotted in figures 25(a) and 25(b). In making left rolls, it was 
impossible for the pilot to apply full aileron and rudder deflection 
simultaneously "because of interference of the pilot 1 s leg movement "by 
the control stick. 

The tests showed that the rudder had enough power to overcome the 
adverse aileron yaw in this condition- with either of the vertical— tail 
configurations . 

Pudder control during take-off and landings .— Both the original and 
production rudders were sufficiently powerful to control the airplane 
during take— off and landing. 



Directional Trim Characteristics 

The rudder trim tab was sufficiently powerful to trim the airplane 
at all speeds above ihO miles per hour. With power on in any configu- 
ration, the rudder force could not "be trimmed to zero below IhO miles 
per hour. The directional trim characteristics with speed are shown 
in figures 26(a) and 26(h) . The rudder force with the original tail 
was high at high speeds and exceeded the 100— pound limit specified in 
the requirements. The rudder force with the production tail also 
exceeded the 100— pound limit of the handling— qualities requirements and 
showed very little improvement over the original tail. 



Aileron Control Characteristics 

The aileron control characteristics of the airplane (figs. 27 to 29) 
were measured by rolling the airplane to the left and to the right by 
abruptly applying various amounts of aileron deflection up to full 
aileron deflection with the rudder held fixed. These tests were 
performed at approximately 5000 feet altitude and at 150, 250, 350, 
and U00 miles per hour calibrated airspeed in the power— on clean condition 
using normal rated power. The variation of rolling acceleration with 
time was always in the correct direction and the rolling velocity varied 
smoothly with 'time at all the speeds tested. These characteristics are 
illustrated in the time histories of figure 27. 

The aileron effectiveness, or helix angle pb/2V and the aileron- 
force variation with total aileron deflection are shown in figure 28. 
These data show the maximum pb/2V obtainable with maximum aileron 
deflection to be 0.06h which occurs at speeds of 150 and 250 miles 
per hour. The aileron forces required to produce the rolls vary smoothly 
with aileron deflection and show no tendency to overbalance. There is 
also enough force to return the controls to neutral when the stick is 
released. The maximum pb/2Y obtainable with a 3^P ourL d stick force 



CONFIDENTIAL 



8 



CONFIDENTIAL 



NACA RM No. L7U1 



is shown in figure 29 and is compared with the Army requirements of 
reference 2. From these data it can "be seen that the pb/2V obtainable 
vith a 30— pound stick force in this airplane is considerably less than 
that specified in the Army requirements. The aileron trim tab vae 
sufficiently powerful to trim the airplane at all the speeds and in all 
the configurations tested. 

CONCLUSIONS 



1. There was no oscillation of the rudder itself during the 
directional oscillations started either from a rudder kick or a sideslip. 
The oscillations produced were always damped to one— half amplitude or 
less in 1 cycle. In the low— speed power-on clean condition, the airplane 
had a rapid spiral divergence when the oscillation was started either 
from a rudder kick or a sideslip with either tail configuration. 

2. With the original vertical tail, the directional stability of 
the airplane in the power— on clean and glide conditions was weak both 
rudder— fixed and rudder— free. The directional stability in the landing, 
approach, and wave— off conditions was satisfactory. The addition of 
the production tail was an improvement but the stability in the power- 
on clean condition was still low. 

3. The effect of higher altitude, approximately 25,000 feet, was 
to decrease slightly the static directional stability. 

k. Tha airplane with the original tail had neutral to negative 
stick— fixed and stick— free dihedral effect at large sideslip angles in 
the low— speed power— on clean condition. This undesirable effect was not 
encountered with the production tail installed on the airplane. 

5. With either the original or production tail the airplane showed 
unsatisfactory characteristics in left aileron rolls with rudder- fixed 
with regard to adverse aileron yaw in the low— speed "power-on clean 
condition. The airplane exhibited rapidly Increasing sideslip angles 
in these maneuvers which would have resulted in dangerous attitudes in 
yaw if the pilot had permitted maximum sideslip angles to be reached. 

6. Both of the rudders tested were sufficiently powerful to 
overcome the adverse aileron. yaw at 150 miles per hour in the power—on 
clean condition and to control the airplane easily during take— off 
and landing. 

7. The rudder— trim— force change with speed was high with either 
tail and the rudder trim tab was not able to trim the airplane with 
power on below ikO miles per hour. 



CONFIDENTIAL 



NACA RM No. L7L11 



CONFIDENTIAL 



9 



• 8. Ths general characteristics of the aileron control were satis— 

•J factory "but the aileron effectiveness was "below the Army requirements. 

.. The maximum aileron effectiveness ph/2V obtainable was 0.064 as 

# * compared to the Army requirement of 0.090. 



• • • 



Langley Memorial Aeronautical Lat oratory 

National Advisory Committee for Aeronautics 
Langley Field, Va. 




Christopher C. Kraft, Jr 
Aeronautical Engineer 



John P. Reeder 
Engineer Test Pilot 





Approved: 

Melvin N. Gough 
Chief of Flight Research Division 

epp 



REFERENCES 



Williams,, Walter C, and Hoover, Herbert H. : Flight Measurements of 
the Directional Stability and Control of a F^51D~NA Airplane 
(AAF No. 4U-13257) with a High-Aspect-Patio Vertical Tail. 
NACA MR No. L5E0^h, Army Air Forces, 19%5. 

Anon.: Stability and Control Characteristics of Airplanes. 
AAF Specifications No. R-1815-A, April 7, 19^5. J&^MtttJ' 



CONFIDENTIAL 



10 



CONFIDENTIAL 



NACA EM No. L7L11 



TABLE I 

PERTINENT DIMENSIONS OF THE P-51H AIRPLANE 



Engine Merlin V-I65O-9 

Propeller (four blades) Aeropro&ucts Model E-20-156-23M5 

Wing area, sq ft 235.73 

Wing span, ft 37.03 

Aspect ratio 5.82 

Wing-flap area, sq ft (two) 31.53 

Aileron area, sq ft (one) 6.35 

Aileron deflection, deg ......... £15 

Total horizontal tail area, sq ft 48.35 

Elevator area, sq ft (one) 6.43 

Original vertical— tail area, sq ft . 23. 40 

Original rudder area, sq ft 10.24 

Rudder deflection of original tail, deg ±30 

Production vertical— tail area, sq ft . . 24.76 

Production rudder area, sq ft 9.77 

Rudder deflection of production tail, deg ±25 

Dorsal fin area, sq ft 1.93 

Original rudder trim tab area, sq ft 0.74 

Unbalancing ratio of original rudder ta"b 0.4:1 

Production rudder trim tat) area, sq ft 0.74 

Unbalancing ratio of production rudder trim ta"b 

Rudder trim tat) deflection, deg 14 left and 8 right 




CONFIDENTIAL 



TABLE II 



VALUES OF DLRECTIONAL— STABILITY PARAMETERS OF P-51H ALRPLANE 
FOR ORIGINAL AND PRODUCTION VERTICAL TAILS 





Speed 


Altitude 


0ri«inal tail 


Production tail 


Condition 


(mph) 


Original 


Production 


d5 r /dP 


dF r /d3 


dSp/dP 


dF r /dp 










(a) 


^aj 


(a) 


(a) 


Glide 


150 


5,000 


5,000 


0.12 


2.0 


0.66 


4.0 




150 


ai a a a 
ci , UUU 


m AAA 
<=*1, UUU 


.14 


n a 

1 . 


.55 


A A 

2.0 




300 


5,000 


4,000 


.51 


20.0 


.70 


25.0 




300 


20,000 


20,000 


.15 


15.0 


.57 


26.0 


Power on clean 


150 


5,000 


7,000 


M 


8.0 


.70 


4.0 




150 


00 c\nn 

d.cL , UUU 


£ij , UUU 


■3k 


Ji n 
t . u 


RR 
• 55 


r n 
P . u 




300 


5,000 


4,000 


.50 


62.0 


.80 


26.0 




300 


22,000 


21,000 


.65 


29.0 


.57 


26.0 




400 


5,000 


5,000 


.31 


43.0 


.84 


39.0 




350 


21,000 


21,000 


.20 


42.0 


.48 


33.0 


Power off landing 


120 


5,000 


5,000 


1.00 


4.0 


1.37 


5.0 




150 


5,000 


5,000 


1.25 


13.0 


1.70 


14.0 


Approach 


120 


5,000 


5,000 


1.00 


6.0 


1.27 


14.0 




150 


5,000 


5,000 


1.20 


14.1 


1.24 


18.0 


Wave-off 


120 


5,000 


5,000 


.83 


9.0 


1.51 


14.0 



o 
o 



a Slopes and taken near zero sideslip. (These values are only approximate 



dp dp 

because of scatter of test points in some conditions tested.) 



MACA^ 



NACA RM No. L7L11 




178.16"- 




CONFIDENTIAL - 133" - 



Figure 1.- Three-view drawing of the P-51H airplane 
with the production tail. 




Figure 2.- Comparison of the original and pro iuction vortical tail assemblies. 

CONFIDENTIAL NACA 




figure Drawing showing comparison or the original and production vertlc 
assemblies. P-^IR airplane. 



al till 




Figure 4.- Three-quarter rear view of P-MH test airplane with production tail. 

CONFIDENTIAL NACA 




Figure 5.- Three-quarter front view of P-blH test airplane with production tail. 

CONFIDENTIAL NACA 



•••• •••• • 




Figure 6.- 



CONF (DENT I AL 



Front view of P-ftlH test airplane. 

NACA 



NACA RM No. L7L11 




Figure 7.- Variation of aileron angle with stick position as measured on 
the ground with no aileron load, P-51H airplane. 



NACA RM No. L7L11 



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Figure 8.- Variation of elevator angle with stick position as measured on the ground 
with no load on the elevator. P-51H airplane* 



NACA RM No. L7L11 



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Figure 9,- Variation of rudder angle with pedal position with no load on the rudder, 
P-51H airplane. This measurement was made with the production tail installed; 
the sane plot aay be used for the original tail by extrapolating the curves to 
30 degrees of right or left rudder. 



NACA RM No. L7L11 




Figure 10.- Variation of c ontrol- friction force with control deflection as measured 
on the ground with no load on the control surfaces. P-51H airplane. 



NACA RM No. L7L11 



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Figure 11.- Decrease in aileron, elevator, and rudder angle due to stretch in the 
control system. P-51H airplane. 



NACA RM No. L7L11 



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fa) 5000 feat altitude. Oscillation started from 
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Figure 12.- Time history of a directional oscillation In the low -speed glide condition. 
P-51H airplane. 



NACA RM No. L7L11 




(b) 50OO feet altitude. Oscillation started from a rudder kick. 
Pilot uaed ailerons to control airplane. Production tail. 



Figure 12.- Continued. 



NACA RM No. L7L11 




fc) -26^ n o feet altitude. Oscillation started from a right 
rudder kick. Production tall. 



Figure 12.- Concluded. 



NACA RM No. L7L11 



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Figure 13.- Time history of a directional oscillation in 
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♦ • 



NACA RM No. L7L11 



• • • 



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(b) 5000 feet altitude. Oscillation started from 
a right rudder kick. Production tail. 



Figure 13.- Continued. 




fc) 20,000 feet altitude. Oscillation started from 
a left rudder kick. Production tail. 



Figure 13.- Concluded. 



NACA RM No. L7L11 




(a> 5°00 feet altitude. Oscillation started 
from a steady sideslip. Original tail. 

Figure ll;.- T<me history of a directional oscillation In the 
low speed power-on clean condition. P-^IH airplane. 



NACA RM No. L7L11 



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?lgure ILl.- Continued. 



• • • 

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NACA RM No. L7L11 



• • • 




(c) 2?,00n feet altitude. Oscillation started from 
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Figure ll+.- Concluded. 



NACA RM No. L7L11 




'a^ 5000 feet altitude. Oscillation started from a steady 

sideslip. Airplane spiralled to left. Original tall. 

Plfrure T<me history of a directional oscillation In the high-apded 

nower-on clean condition. P-^IH airplane. 



NACA RM No. L7L11 




(b) 10,000 feet altitude. Oscillation started 
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Figure 1^.- Continued. 



NACA RM No. L7L11 



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Figure 1^.- Concluded. 



NACA RM No. L7L11 




(ml 5^°0 feet altitude. Oscillation started by a left rudder kick. 
Maneuver re rf ormed in rough air. Original tail. 

Figure 16.- Time history of a directional oscillation in the landing condition. P-51H airplane. 



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Plffure 1 6.- Concluded. 



NACA RM No. L7L11 




(a) 1 52- ml les-per-hour, calibrated airspeed at $000 feet 
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Figure 17.- Directional stability and control characteristics of the 
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NACA RM No. L7L11 





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Figure 1?.- Concluded, 



NACA RM No. L7L11 




'a^ lS'i -miles-per-hour, calibrated airspeed at 21,000 feet 
average altitude. Original tall. 



PI cure 19.- Directional stability and control characteristics of the P-5IH 
airplane In the power-off clean condition at high altitude. 



NACA RM No. L7L11 




(b) l^U-miles-per-hour calibrated airspeed at 21,000 feet 
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Figure 18.- Continued. 



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average altitude. Original tall. 



figure 1^.- Continued. 



NACA RM No. L7L11 



CONFIDENTIAL 




fd) 502-mile a -per-h our calibrated airspeed at 20,000 feet 
average altitude. Production tall. 



Figure 13.- Concluded, 



NACA 



RM No. L7L11 




ttk) 153-rll es-rer-hour, calibrated a'r?"-*ed at ^f\jTeet 
average altitude. Original tall. 



Figure 19.- Directional stability and control characteristics of the 
P-51H airplane In the power-on clean condition at low altitude. 



NACA RM No. L7L11 




fb) l^l miles rer- hour calibrated airspeed at 7000 feet 
average altitude. Production tail. 



Plgure 13.- Continued. 



NACA RM No. L7L11 




(c) 506"i11 ea- rer-hour calibrated air- 
SDeed at $000 feet average altitude. 
Original tail. 



(d) UoU-mil es-per-hour calibrated air- 
greed at 5000 feet average altitude. 
Original tail. 



Figure 19.- Continued. 



NACA RM No. L7L11 




(e) 500- relies- per- hour calibrated air- 
speed at l|0OO feet average altitude. 
Production tall. 



(f) I4.05-1T1I les-rer-hour calibrated air- 
speed at ^000 feet average alt1tv.de. 
Production tall. 



Figure 19.- Concluded. 



NACA RM No. L7L11 




Figure 20.- Directional stability and control characteristics of the P-^IH 
airplane in the power-on clean condition at high altitude. 



NACA RM No. L7L11 




(b) miles per -hour calibrated airspeed at 25,000 feet 

average altitude. Production tall. 

Figure 20.- ConMnue'd. 



NACA RM No. L7L11 




(c) 50U-mll«s -per-hour calibrated air- 
speed at 22,000 feet average altitude. 
Original tall. 



(d) 255- mllea -per-hour calibrated air- 
speed at 21,000 feet average altitude. 
Original tall. 



Plgure 20.- Continued. 





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Figure 21.- Directional stability and control characteristics of the 
P-51H airplane In the power-off landing condition. 



NACA RM No. L7L11 



CONFIDENTIAL . 




(b) 



ll' 7 miles-rer-hour calibrated airsreed at ^COO feet 
average altitude. Production tail. 



Figure 21.- Continued. 



NACA RM No. L7L11 




(c) 15^" wile 8- per- hour calibrated airspeed at ^OOO feet 
average altitude. Original tall. 



Figure 21.- Continued. 



NACA RM No. L7L11 



CONFIDENT I AL 




(d) 153-Trlles-rer-hour calibrated airspeed at 5000 feet 
average altitude. Production tall. 

Figure 21.- Concluded. 



NACA RM No. L7L11 




fa^ 120- rrlles -per- hour calibrated airsneed at 50°° feet 
average altitude. Original tall. 



PI pure 22.- Dl recti onal stability and control characteristics of the 
P-51H airplane In the approach condition. 



NACA RM No. L7L11 




fb) 122-i".ile5-rer-hour calibrated airspeed at ^COO feet 
average altitude. Production tail. 

Figure 22.- Continued. 



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Figure 22.- Continued. 



NACA RM No. L7L11 




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Figure 22.- Concluded. 



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Plgure 25.- Directional stability and control cbaracterl "t 1 cs of the 
P-51R airplane In the wave-off condition. 



NACA RM No. L7L11 









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Pigure 23.- Concluded. 



NACA RM No. L7L11 











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Figure 24.- Maximum change in sideslip attained during abrupt aileron rolls out of 45 degree 
banked turns with rudder fixed. Different symbols indicate different flights. P-51 H 
airplane. CONFIDENTIAL 



NACA RM No. L7L11 




(b) Production tail. 
Figure 24#* Concluded. 



NACA KM No. L7L11 



























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Figure 25.- Maximum change in sideslip attained during abrupt aileron rolls out of U5 degree 
banked turns using full aileron deflection and varying amounts of rudder deflection. 
152- miles -per -hour calibrated airspeed in the power-on clean condition. P-51H airplane. 



CONFIDENTIAL 



NACA KM No. L7L11 



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Figure 25.- Concluded. 



NACA RM No. L7L11 




(a) Original tail. 



Figure 26.- Directional trirr. characteristics of the P-^IH airplane In 
the power-on clean condition. 



NACA RM No. L7L11 



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PI gure 26.- Concluded. 



NACA RM No. L7L11 




fa) Ri?ht roll. 



fb) Left roll. 



Pigure 27.- Time history of tyrlcal aileron rolls performed with rudder held fixed 
to measure the aileron effectiveness pb/2V. F-^IH airplane. 



CONFIDENTIAL . 




Figure 2R.- Variation of helix angle. pb/2V with total aileron deflection 
In the power-on clean condition at an average altitude of ^000 feet. 
Normal rated power. T-^IH airplane. 




Plgure 2^.- Concluded. 



NACA 



RM No. L7L11 





























































































































































































































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Figure 20,- Maximum pb/2V and rolling velocity obtainable with a 50-pound stick 
force as a function of speed. These tests were made on the production air- 
plane at an average altitude of ^000 feet. F-51 H airplane.